This invention relates to a gas turbine engine and, more particularly, to the prevention of wear damage between the rotor blades and the rotor disk in the compressor and fan sections of the engine.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is combusted, and the resulting hot combustion gases are passed through a turbine mounted on the same shaft. The flow of gas turns the turbine by contacting an airfoil portion of the turbine blade, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. There may additionally be a bypass fan that forces air around the center core of the engine, driven by a shaft extending from the turbine section.
The compressor and the bypass fan are both rotating structures in which blades extend radially outwardly from a rotor disk. In most cases, the blades are made of a different material than the rotor disk, so that they are manufactured separately and then affixed to the rotor disk. That is, compressor blades are manufactured and mounted to a compressor rotor disk, and fan blades are manufactured and mounted to a fan rotor disk.
In one approach that is widely used, each blade has an airfoil-shaped region and a root at one end thereof. The root is in the form of a dovetail structure. The rotor disk has corresponding hub slots therein. The dovetail structure of each root slides into its respective hub slot to affix the blade to the rotor disk.
When the gas turbine engine is operated, there is a high-frequency, low amplitude relative movement between the root and the surface of the hub slot. This movement produces wear damage, of a type typically termed xe2x80x9cfretting wearxe2x80x9d, to the root or to the hub slot. The fretting wear may lead to the initiation of fatigue cracks which in turn lead to the need for premature inspections of the components, or in extreme cases may lead to failure.
This problem has long been a concern to aircraft engine manufacturers. A variety of anti-wear coatings have been developed. However, these coatings have not been entirely satisfactory for compressor and fan rotor applications. There is a need for a more suitable protective coatings. The present invention fulfills this need, and further provides related advantages.
The present invention includes a method for providing a rotating structure of a gas turbine engine. The contact between the rotor disk and the rotor blades is protected by a protective coating that reduces friction and wear between these components. The result is an extended life without wear-based fatigue damage and failures.
A method for providing a rotating structure of a gas turbine engine comprises the steps of furnishing a rotor disk comprising a hub with a plurality of hub slots in a periphery of the hub. Each hub slot has a hub slot surface. A plurality of rotor blades are furnished, wherein each rotor blade comprises an airfoil, and a root at one end of the airfoil. The root is shaped and sized to be received in one of the hub slots of the rotor disk. A protective coating is deposited at a location which will be, upon assembly, disposed between the root of each rotor blade and the respective hub slot surface. The deposition is performed by a wire arc spray process, preferably a compressed-air wire arc spray process. The protective coating is a protective alloy comprising (preferably consisting essentially of), in weight percent, from about 6.0 to about 8.5 percent aluminum, from 0 to about 0.5 percent manganese, from 0 to about 0.2 percent zinc, from 0 to about 0.1 percent silicon, from 0 to about 0.1 percent iron, from 0 to about 0.02 percent lead, remainder copper and impurities. The protective coating is preferably from about 0.003 to about 0.020 inch thick. The roots of the rotor blades are assembled into the respective hub slots of the rotor disk to form the rotating structure.
The rotor disk may be a compressor disk, and the rotor blades are compressor blades. Alternatively, the rotor disk may be a fan disk, and the rotor blades are fan blades. Preferably, the hub of the rotor disk is made of a titanium alloy.
The protective coating may be deposited on the root, or on the hub slot surface, or both. Alternatively, the protective coating may be deposited on a shim that is subsequently positioned during assembly between the root and the hub slot surface.
The rotating structure is thereafter operated such that the root is at a temperature of from about 75xc2x0 F. to about 350xc2x0 F.
In a preferred form, a method for providing a rotating structure of a gas turbine engine comprises the steps of furnishing a set of rotor blades, with each rotor blade comprising an airfoil, and a root at one end of the airfoil. A protective coating having the protective alloy composition set forth above is deposited on the root of each rotor blade by a wire arc spray process. The rotor blades are assembled into the hub slots of the rotor disk and subsequently operated.
The present approach yields a low-friction, low-wear interface between the root of the blade and the hub slot surface of the rotor disk. The wire arc spray process produces good bonding between the protective coating and the substrate, with a relatively low-temperature deposition technique that does not overly heat the substrate or produce high differential thermal stresses between the substrate and the protective coating. The preferred compressed-air wire arc spray process has the additional advantage that no contaminants such as hydrocarbons are introduced into the deposited protective coating.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.